Method for designing flowfield molded hypersonic inlet for integrated turbojet and ram-scramjet applications

ABSTRACT

A diverterless hypersonic inlet (DHI) for a high speed, air-breathing propulsion system reduces the ingested boundary layer flow, drag, and weight, and maintains a high capture area for hypersonic applications. The design enables high vehicle fineness ratios, low-observable features, and enhances ramjet operability limits. The DHI is optimized for a particular design flight Mach number. A forebody segment generates and focuses a system of multiple upstream shock waves at desired strengths and angles to facilitate required inlet and engine airflow conditions. The forebody contour diverts boundary layer flow to the inlet sides, effectively reducing the thickness of the boundary layer that is ingested by the inlet, while maintaining the capture area required by the hypersonic propulsion system. The cowl assembly is shaped to integrate with the forebody shock system and the thinned boundary layer region.

This patent application is subject to U.S. Air Force contract No.F33615-00-D-3053. The Government has certain rights in this invention.

BACKGROUND OF THE INVENTION

1. Technical Field

The present invention relates in general to air craft engines and, inparticular, to an improved system, method, and apparatus for adiverterless hypersonic inlet for integrated turbojet and ram-scramjetaircraft jet engine applications.

2. Description of the Related Art

Boundary layer diverters using splitter plate and wedge geometries aretraditionally employed to reduce or eliminate boundary layer thicknessupstream from the inlets of airbreathing propulsion systems at Mach 1 toMach 2+. The high drag and weight of such devices has been overcome inthe past with more thrust (which required bigger engines), more engines,or the additional use of afterburners. These designs result in yet moreweight, less vehicle payload capacity, higher fuel consumption, and/orother aircraft design penalties. Higher Mach number, multi-engineaircraft (Mach 2 to 3+) have employed efficient, axisymmetric spikeinlets or underwing-mounted nacelles. However, these designs requirelarge amounts of internal boundary layer bleed, add significant drag,and are structurally inefficient.

The development of the ramjet or scramjet powered hypersonic aircrafthas been historically thwarted by inefficient inlet and forebodydesigns. An inefficient hypersonic forebody and inlet design can havepoor transonic performance characteristics, which compromises theability of an on-board gas turbine propulsion system to accelerate toramjet transition speed without running out of fuel. An inefficientdesign can also allow ramjet ingestion of a thick boundary layer whichdelays ramjet startup until the vehicle achieves at least Mach 3.5. Gasturbine propulsion systems are historically inefficient beyond Mach 2.5,so afterburning, base burning, rocket assistance, releasing the vehiclefrom an aircraft at altitude, and combinations thereof are generallyconsidered to fill the gap between Mach 2.5 and 3.5, and thereby achieverequired ramjet start-up conditions. As such, the problems associatedwith hypersonic inlets (boundary layer flow, high transonic drag, etc.)have not been resolved to facilitate the practical use of efficient,self-powered, hypersonic airbreathing vehicles. An improved solutionthat addresses these limitations would be desirable.

SUMMARY OF THE INVENTION

One embodiment of a system, method, and apparatus for diverterlesshypersonic inlets (DHI) offers solutions that incorporate favorable airflow characteristics for high speed (e.g., Mach 3+) air-breathingpropulsion systems. The DHI shape thins the boundary layer for enginesthat traditionally operate above Mach 3, thereby allowing them to workat lower Mach numbers. The long and narrow shape of DHI designs alsoimproves transonic drag, which is a common issue withtraditionally-shaped (i.e., wide and flat) hypersonic vehicles. The DHIreduces the penalties associated with traditional boundary layerdiversion techniques and enables vehicle operability well into thehypersonic flight regime.

The present invention is well suited for ramjet and/or scramjet engines.The DHI reduces the ingested boundary layer flow, reduces drag andweight associated with traditional boundary layer diverterconfigurations, maintains a high capture area for hypersonicapplications, enables high vehicle fineness ratios (which lowerstransonic drag), provides a practical basis for low-observable features,and enhances ramjet operability limits.

In one embodiment, the DHI comprises a cowl assembly mounted to asegmented, contoured forebody. The DHI is designed to fit vehicle sizeconstraints and is optimized for specific flight Mach numbers. Theforebody includes segments that generate and focus a system of multipleupstream shock waves at desired strengths and angles to facilitaterequired inlet and engine airflow conditions. The forebody contourdiverts boundary layer flow to either side of the inlet, effectivelyreducing the thickness of the boundary layer that is ingested by theinlet, while maintaining the capture area required by a hypersonicpropulsion system. The cowl assembly is shaped to integrate with theforebody shock system and the thinned boundary layer region. The outermold line of the DHI enables the vehicle to maintain a high finenessratio which is beneficial for transonic acceleration.

The DHI has been computationally proven for at least Mach 3 to Mach 10applications, which is well above the nominal operating envelope oftraditional boundary layer diverting systems. The DHI is highlyeffective in boundary layer reduction above Mach 3, which is importantfor ramjet and scramjet applications. The present invention facilitatessuccessful engine start-up and enhanced operability and performance atmuch lower Mach numbers than those demonstrated in the prior art. Alower Mach number start speed for a ramjet system closes the gap betweenthe maximum speed of a gas turbine accelerator and the minimum speed atwhich a ramjet can take over. This design enables a dual mode, airbreathing vehicle to propel itself to hypersonic speeds from a standingstart on the ground. The DHI utilizes an elegant design method thatconverges on an optimized solution with few iterations, minimizingcomputer time, and requiring relatively few man hours. It provides auniform flowfield at the cowl plane and can be modified to includelow-observable features.

The foregoing and other objects and advantages of the present inventionwill be apparent to those skilled in the art, in view of the followingdetailed description of the present invention, taken in conjunction withthe appended claims and the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

So that the manner in which the features and advantages of theinvention, as well as others which will become apparent are attained andcan be understood in more detail, more particular description of theinvention briefly summarized above may be had by reference to theembodiment thereof which is illustrated in the appended drawings, whichdrawings form a part of this specification. It is to be noted, however,that the drawings illustrate only an embodiment of the invention andtherefore are not to be considered limiting of its scope as theinvention may admit to other equally effective embodiments.

FIG. 1 is a plot of engine performance versus speed for turbojet,ramjet, and scramjet engines;

FIG. 2 is an isometric view of one embodiment of an aircraft having adiverterless hypersonic inlet (DHI) constructed in accordance with thepresent invention;

FIG. 3 is a front isometric view rendering comparing a DHI designaugmented with and without boundary layer control;

FIG. 4 is a side view rendering comparing boundary layer thickness forthe DHI and a two-dimensional wedge-shaped inlet;

FIG. 5 is a side view rendering of the shock system for one embodimentof a DHI design constructed in accordance with the present invention;

FIG. 6 illustrates isometric view renderings of a flowfield generator(FFG) spanwise contour for one embodiment of a DHI design constructed inaccordance with the present invention;

FIG. 7 illustrates isometric view renderings of the FFG of FIG. 6 withstreamline seeds and planes for one embodiment of a DHI designconstructed in accordance with the present invention; and

FIG. 8 is a high level flow diagram of one embodiment of a methodconstructed in accordance with the present invention.

DETAILED DESCRIPTION OF THE INVENTION

Referring to FIGS. 1-8, various embodiments of a system, method, andapparatus for a diverterless hypersonic inlet (DHI) constructed inaccordance with the present invention are shown. As shown in FIG. 1, theperformance 11 of aircraft with conventional turbojet enginesdeteriorates quickly as flight speed approaches Mach 3. Similarly, theperformance 13 of conventionally configured ramjets and scramjets isalso weak at flight speeds in the vicinity of Mach 3. However, a highspeed aircraft configured with a DHI constructed in accordance with thepresent invention has a performance 15 with a significantly enhancedregion of improvement 17 over conventional inlet designs. The turbojet,ramjet, and scramjet engine applications for the DHI of the presentinvention include flight speeds of approximately Mach 2.5 to 10.

FIG. 2 illustrates an aircraft 21 with one embodiment of a DHI design 23for processing hypersonic airflow. The DHI 23 extends along the forebodyand incorporates the cowl 27 to define an intake manifold for theaircraft's engine. The forebody has a longitudinal axis 31 extendingalong a length of the aircraft 21, a lateral axis 33 orthogonal to thelongitudinal axis 31, and a transverse axis 35 orthogonal to both thelongitudinal and lateral axes 31, 33. In one embodiment, the DHI 23 alsoconfigures the forebody with a high outwash angle 37 (e.g., at least3.9°) with respect to the longitudinal axis 31 for spilling the boundarylayer airflow.

As best shown in FIG. 5, the DHI 23 includes a plurality of rampsegments 41, 43, 45 (three shown) that are formed in a profile of theforebody extending along the longitudinal axis 31. Although theillustrated embodiment depicts three ramp segments 41, 43, 45, more orfewer segments may be utilized. Each of the ramp segments 41, 43, 45 areformed at angles with respect to the longitudinal and transverse axes31, 35, such that the ramp segments 41, 43, 45 form a series of anglesthat sequentially increase in a downstream direction toward the cowl 27.As a result, the forebody of DHI 23 has significantly greater length 61(lower half of FIG. 4) than the length 63 of a conventionally equippedinlet (upper half of FIG. 4). Even though the DHI is longer thanconventionally equipped inlets, the boundary layer is proportionallythinner, as explained below.

The DHI 23 also comprises spanwise contours 47 that are formed on theforebody adjacent each lateral side of the cowl 27. The spanwisecontours 47 extend along both the longitudinal and lateral axes 31, 33.As best shown in FIG. 6, the spanwise contours 47 are designed toeffectively divert and thin the boundary layer airflow for turbojet,ramjet, and scramjet engine applications. In contrast, an aircraft 51with a conventional inlet 53 (top of FIG. 4) develops excessive boundarylayer conditions 55 that degrade the performance 13 (see FIG. 1) of highspeed aircraft engines. In one embodiment, the spanwise contours 47 aresuper elliptical as defined by the equation: (x/a)^(n)+(y/b)^(n)=1,where n>2 and a and b are the ellipse radii.

In addition, the DHI 23 enables the aircraft 21 to have a fineness ratioin excess of six (6). For purposes of the present invention, finenessratio is defined as the length of the aircraft divided by the equivalentcircular diameter of the maximum frontal cross-sectional area of theaircraft.

Referring now to FIG. 8, the present invention also comprises a methodof designing a hypersonic inlet for an aircraft. In one embodiment, themethod begins as indicated at step 81, and comprises establishing acenterline geometry for a flowfield generator (FFG) (step 83);optimizing the FFG and a shock system 121 (FIG. 5) configuration to fitaircraft geometrical constraints and engine airflow requirements (step85); establishing spanwise contours for the FFG (step 87); gridding andrunning computation fluid dynamics (CFD) for the FFG (step 89); anditerating steps (a) through (d) until the desired cowl plane Machnumber, flowfield uniformity, and geometrical constraints are achieved(step 91).

In one embodiment, the method further comprises streamline tracing 123(FIG. 7) a diverterless hypersonic inlet (DHI) forebody within athree-dimensional, FFG CFD solution and transferring the streamlinesinto a computer aided design (CAD) environment (step 93); surfacingstreamlines of the DHI forebody in the CAD environment and consolidatingthem with a cowl assembly (step 95); gridding DHI geometry and runningviscous CFD analysis (step 97); modifying the DHI geometry andconfirming with follow-on viscous CFD runs (step 99); and thenintegrating the DHI geometry with geometry of the aircraft foraerodynamic analysis (step 101), before ending as indicated at step 103.

In step 83, the method also may comprise performing a series of inviscidconical and two-dimensional shock calculations to determine viable shocksystem configurations, including a plurality of shocks and shock angles,to achieve the desired cowl plane Mach number; and designing asegmented, cone-shaped FFG required to generate the desired shock systemconfiguration. In step 85, the method may comprise factoring an intendedstreamline cut position of approximately 10% to 40% of the cowl height.

In step 87, the method may comprise contouring each spanwise contourindividually, wherein each spanwise contour affects a spanwise shape ofgenerated shocks, lateral boundary layer migration, and capture area ofthe cowl. In step 89, the method may comprise constructing streamlineseed planes, superimposing construction planes to represent aircraftconstraints and cowl geometry, assessing cowl plane flowfield uniformityand Mach number, assessing shock positions with respect to cowlgeometry, and running streamlines forward from the cowl plane to assessa capture area thereof.

In addition, step 97 may comprise reassessing shock positions, cowlplane Mach number, and capture area, and assessing boundary layerthickness and surface pressure gradients; and step 99 may compriseevaluating off-design conditions and augmenting with DHI forebody flowcontrol.

While the invention has been shown or described in only some of itsforms, it should be apparent to those skilled in the art that it is notso limited, but is susceptible to various changes without departing fromthe scope of the invention.

1. A method of designing a hypersonic inlet for an aircraft, the methodcomprising: (a) establishing a centerline geometry for a flowfieldgenerator (FFG); (b) establishing spanwise contours for the FFG; (c)gridding and running computational fluid dynamics (CFD) for the FFG; (d)iterating steps (a) through (c) until a desired inlet aperture flowcharacteristics uniformity, and geometrical constraints are achieved;(e) streamline tracing a diverterless hypersonic inlet (DHI) forebodywithin a three-dimensional, FFG CFD solution and transferring thestreamlines into a computer aided design (CAD) environment; and (f)surfacing streamlines of the DHI forebody in the CAD environment andconsolidating them with an inlet aperture design.
 2. A method accordingto claim 1, wherein step (a) comprises performing a series of shockcalculations to determine viable shock system configurations, includingat least one of a shock and a shock angle, to achieve the desired inletaperture flow characteristics; and based on the shock calculations,designing an FFG required to generate a desired shock systemconfiguration.
 3. A method according to claim 1, further comprisingfactoring an intended streamline cut position of approximately 10% to40% of the inlet aperture.
 4. A method according to claim 1, whereinstep (b) comprises contouring each spanwise contour individually,wherein each spanwise contour affects a spanwise shape of generatedshocks, lateral boundary layer migration, and a capture area of theinlet aperture.
 5. A method according to claim 1, wherein step (c)comprises constructing streamline seed planes, applying aircraftconstraints and inlet aperture geometry, assessing inlet apertureflowfield characteristics, assessing shock positions, and calculatingstreamlines forward from the inlet aperture to assess a capture areathereof.
 6. A method according to claim 1, further comprisingreassessing shock positions, inlet aperture flow characteristics, andcapture area, and assessing boundary layer thickness and surfacepressure gradients.
 7. A method according to claim 1, further comprisingevaluating off-design conditions and augmenting with DHI forebody flowcontrol.
 8. A method of designing a hypersonic inlet for an aircraft,the method comprising: (a) establishing a centerline geometry for aflowfield generator (FFG); (b) optimizing the FFG and a shock systemconfiguration to fit aircraft geometrical constraints and engine airflowrequirements; (c) establishing spanwise contours for the FFG, andcontouring each spanwise contour individually, wherein each spanwisecontour affects a spanwise shape of generated shocks, lateral boundarylayer migration, and a capture area of a cowl; (d) gridding and runningcomputation fluid dynamics (CFD) for the FFG by constructing streamlineseed planes, applying aircraft constraints and inlet aperture geometry,assessing inlet aperture flowfield characteristics, assessing shockpositions, and calculating streamlines forward from the inlet apertureto assess a capture area thereof; (e) iterating steps (a) through (d)until a desired inlet aperture flow characteristics and geometricalconstraints are achieved; (f) streamline tracing a diverterlesshypersonic inlet (DHI) forebody within a three-dimensional, FFG CFDsolution and transferring the streamlines into a computer aided design(CAD) environment; and (g) surfacing streamlines of the DHI forebody inthe CAD environment and consolidating them with an inlet aperture design9. A method according to claim 8, wherein step (a) comprises performinga series of shock calculations to determine viable shock systemconfigurations, including a at least one of a shock and shock angle, toachieve the desired inlet aperture flow characteristics; and designingan FFG based onthe shock calculations required to generate a desiredshock system configuration.
 10. A method according to claim 8, whereinstep (b) comprises factoring an intended streamline cut position ofapproximately 10% to 40% of the inlet aperture.
 11. A method accordingto claim 8, further comprising reassessing shock positions, inletflowfield characteristics, and capture area, and assessing boundarylayer thickness and surface pressure gradients.
 12. The method accordingto claim 8, further comprising: gridding DHI geometry and runningviscous CFD analysis; modifying the DHI geometry and confirming withfollow-on viscous CFD runs; evaluating off-design conditions andaugmenting with the DHI with forebody flow control; and integrating theDHI geometry with geometry of the aircraft for aerodynamic analysis.